In spite of the abundance and sophistication of radio navigation systems available today, it is not unusual for an aircraft during a long flight mission to be occasionally cut off from radio navigation signals for significant periods. Although a self-contained inertial navigation system would alleviate such a problem, such systems are just too expensive for many aircraft owners. Known inertial navigation systems capable of an accuracy of one nautical mile per hour cannot be implemented at prices low enough to attract general aviation customers. Present day inertial navigation systems of the 1 nm/hr class cost in the $50,000--$100,000 range.
Conventional inertial navigation systems are based on measurement and double integration of the linear accelerations of the vehicle. In one class of systems a plurality of orthogonally disposed accelerometers are mounted on a gyroscopically stabilized platform that is supported by three, four or even five gimbals. A key element in the development of this inertial system has been the stabilized platform on which the accelerometers are mounted and whose attitude is controlled by the gyroscopes. The platform performs two critical functions. It establishes a coordinate system for the accelerometers and isolates them from angular motions of the vehicle. The first function simplifies navigation computations while the second simplifies computations as well as sensor design. The accelerometers must be precisely horizontally aligned. If they have any tilt, they record acceleration due to gravity and this is added to the measured horizontal acceleration. As the effect of gravity is continuous and provides an acceleration that is usually larger than that experienced by the aircraft, accelerometer misalignment may cause considerable errors. For example, if an accelerometer sensed 0.1% of the gravity field, double integration of that acceleration would produce approximately an 18-mile position error after a minute of flight. A representative gimballed inertial system is described in "Guidance System," U.S. Pat. No. 3,104,545, C. S. Draper et al., Sept. 24, 1963.
In another class of systems, known as strapdown inertial navigation systems, a plurality of orthogonally displaced gyros and a plurality of orthogonally aligned accelerometers are strapped directly to the airframe. In the strapdown inertial system the stable platform is replaced by two computer functions: one to establish a coordinate reference based on the gyro outputs and the other to transform the accelerometer outputs into the established coordinate frame. The strapdown sensor must operate to the same order of accuracy as its platform counterpart. The angular rate regime for the strapdown sensor is in deg/sec instead of the deg/hr category for a stable platform sensor. The major design impact is of course on the gyro rather than the accelerometer, as the former must measure the rates very precisely. The dynamic range for the gyro could, for example, begin at a threshold of 0.01 deg/hr and extend to the area of 100 deg/sec or so, for a span of about 10.sup.7 compared with a 10.sup.4 span in gimballed platforms.
Efforts to improve reliability and to reduce the cost and size of inertial navigation systems have, in the past decade, been largely directed to strapdown systems. "Strapdown Navigation Technology: A Literature Survey," Garg et al., AIAA Journal of Guidance and Control, Vol. 1, No. 3, May-June 1978, pp. 161-172. In strapdown systems the advantages gained from dispensing with gimbals are largely off-set by the increased demands placed on the gyroscopes and the computational apparatus. The greater sensitivity of strapdown systems to alignment errors, and to correlated noise components in the accelerometers and gyroscopes, have so far impeded their successful competition with gimballed systems in the 1 nm/hr class. "Advantages of Gimballed Inertial Navigation Systems," Roland Peterson, Proceedings of the IEEE 1976 National Aerospace and Electronics Conference, pp. 508-514.
A few departures from the aforementioned types of inertial systems have been noted. The article "Kalman Filter Divergence and Aircraft Motion Estimators," A. E. Bryson, AIAA Journal of Guidance and Control, Vol. 1, No. 1, Jan.-Feb. 1978, pp. 71-79, describes a method of estimating velocities from measurements of roll rate, pitch rate, heading and altitude. Although it may have possibilities for short-term navigation, further work is needed in the areas of parameter variation sensitivity, trim error and wind modeling. Hector ("The RAMP Inertial Navigation System," Philips Technical Review, Vol. 29, 1968, No. 3/4, pp. 69-85) discloses an inertial system in which two single-axis pendulums are supported and orthogonally aligned on a stabilized platform. Displacement of the vehicle is determined by twice integrating the angular acceleration of each pendulum. Monaco et al., ("Schuler Tuned Vertical Indicating System," AIAA Journal of Guidance and Control, Vol. 1, No. 6, Nov.-Dec. 1978, pp. 413-419) show a gyro-less, vertically-oriented, single-axis pendulum that they say can be employed in a navigation system. The motion of the single-axis pendulum is sensed by an angular accelerometer. The sensor output is amplified and fed to a torque generator that positions the pendulum. Both Hector and Monaco suggest that their pendulums be Schuler tuned by artificially increasing the moment of inertia by a factor of the order of 10.sup.8. Such a scheme calls for each torque generator to be preceeded by an amplifier with a gain of approximately 10.sup.7. The problems attendant with such a massive gain (stability, noise, etc.) raise doubts as to how such systems would perform in an actual flight environment. For example, if the amplifier input noise or gyro pick-off noise were only 10 .mu.v, the amplifier would easily be saturated and would induce a large extraneous torque in the pendulum.